1. Field of the Invention
The invention relates generally to the field of fabricating composite structures. More specifically, the invention relates to the field of manufacturing composite structures for aircraft.
2. Description of the Related Art
The disclosed embodiments provide systems and methods for fabricating with composite material the fuselage of an aircraft, among other structures. The fuselage of an aircraft is the hollow body of an aircraft, and may comprise a skin. Structural members such as stringers and frames may be fastened or otherwise secured to the inner surface of the skin to provide structural integrity to the fuselage. Numerous aircraft parts, such as the wings, the landing gear, et cetera, may be secured to the fuselage. Proper and efficient fabrication of the fuselage is therefore of prime importance in the construction of an aircraft.
The fuselage of an aircraft, including its structural supporting members, is customarily made of aluminum or other metals. A fuselage made from metal, however, is generally heavy and excessively thick, and may cause unwanted drag during flight. More recently, thus, many aircraft manufacturers are fabricating the entire fuselage or some of its components with composite material.
FIG. 1 shows a section of a fuselage section 10 made of composite material. The fuselage 10 has a curved skin 12 to which stringers 14 and frames 16 are bonded or otherwise secured. Securement of the stringers 14, frames 16 (and other structural members of the fuselage 10) to the skin 12 has heretofore been a laborious process, and has generally been effectuated in one of three ways. Specifically, the skin 12, the stringers 14, and the frames 16 may be independently fabricated with composite material. Each stringer 14 and frame 16 may then be bonded to the skin 12 piece by piece. As can be appreciated, this process may require individual tooling for the securement of each component 14, 16, to the skin 12, which may be fairly time consuming. Alternatively, bond mandrels may be used to co-cure some of the components 14, 16 to the skin 12. The mandrels may have to be removed after curing, and other parts may have be cured to the skin 12 at locations previously occupied by the mandrels. Or, extensive positional tooling may be used to co-cure the components 14, 16, and the skin 12 all at once. Configuration of such extensive tooling may be an arduous task and require much labor.
U.S. Pat. No. 5,242,523 to Willden et al., discloses a method whereby uncured stiffeners and pre-cured frame members are laid on the surface of the skin. A flexible caul is placed between every set of two frame members, and an elongated mandrel is placed under each stiffener. The cauls define the outer shape of the structure, and the mandrels define the inner shape of the structure. The entire structure is placed within a vacuum bag. During this co-curing, the bag presses the frame members against the skin panels, and presses the cauls causing them to press the stiffeners against the skin panel, thus forming the finished structure.
U.S. Pat. No. 8,220,154 to Cacciaguerra discloses the use of “shell tooling” for fabricating the fuselage. The frames and structural parts are assembled around a chuck resting on a frame structure. Shell fixtures are then installed in spaces between the frames, and fastened to the frames by detachable fasteners. The shell fixtures have to be removed after the skin is coupled to the framework.